This invention relates to gas turbine engines and more particularly to a flow directing element such as a stator vane or a blade for the compressor or turbine section of such engines.
Modern gas turbine engines employ axial flow compressors or centrifugal compressors to pressurize air to a pressure and temperature suitable for combustion. The compressor directs air from a region of low pressure at the compressor intake to a region of higher pressure at the compressor discharge. Because air naturally flows from a region of high pressure to a region of low pressure, the compression process involves considerable expenditure of energy, and it is desirable to minimize this expenditure by making the compressor as efficient as possible.
Because the direction of airflow through the engine compressor is opposite to the natural tendency of air to flow in the direction of decreasing pressure, compressors are susceptible to an aerodynamic instability referred to as surge. Compressor surge is a violent event involving the sudden and unanticipated reversal of the airflow direction through the compressor. At best, a surge results in the momentary loss of engine power followed by a resumption of the normal airflow through the compressor and normal operation of the engine. In more severe cases, normal airflow through the compressor is not readily reestablished, resulting in a sustained loss of engine power. Moreover, the violent character of a surge can damage engine components.
Clearly, surge is a phenomenon to be avoided. To avoid surge, compressor designers seek to increase the surge margin of a compressor. As explained more fully hereinafter, surge margin is a measure of a compressor""s resistance to surge. Unfortunately, and as is well known in the art, increasing a compressor""s surge margin results in diminished compressor efficiency and increased fuel consumption during sustained, steady state operation of the engine.
What is needed is a means for improving compressor surge margin without compromising compressor and engine efficiency.
In accordance with the present invention, the airfoil portion of a blade or vane for a turbine engine is uniquely shaped to redistribute the airflow velocity across the span of the airfoil. In particular, the airfoil has a chord length which increases from a minimum value at the airfoil root to a greater value at a part span location intermediate the root and the airfoil tip. Between the part span location and the tip, the chord length is substantially constant.
In one embodiment of the invention, the part span location is preferably between 25% and 75% of the airfoil span. In another embodiment the ratio of the chord length at the airfoil root to the chord length at the part span location is between 0.7 and 0.9.
The invention has been demonstrated to be especially effective when used as a stator vane in at least some of the fixed stages of an axial flow compressor for an aircraft gas turbine engine.
The primary advantage of the invention is its unexpected ability to increase the compressor""s surge margin while not degrading its steady state efficiency.
These and other advantages and features of the invention will become more apparent in view of the following discussion of the best mode for carrying out the invention and the accompanying drawings.